Gas turbine engine with improved dynamic characteristics

ABSTRACT

A gas turbine engine for an aircraft includes a fan system including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode including a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A forward whirl frequency margin of: 
     
       
         
           
             
               
                 
                   
                     
                       
                         
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     is in the range from 10 to 100%.

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUK Patent Application Number 2008334.1 filed on Jun. 3, 2020, the entirecontents of which are incorporated herein by reference.

BACKGROUND 1. Field of the Disclosure

The present disclosure relates to aircraft engines with improved dynamiccharacteristics, and more specifically to engines with improved handlingof vibrational modes by the avoidance of frequency coincidence betweennatural frequencies and their potential excitation sources, and methodsof using such engines.

SUMMARY

According to a first aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor;and a fan system comprising: a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a fan shaft. Theengine further comprises a gearbox, and a gearbox output shaft arrangedto couple an output of the gearbox to the fan shaft. The gearboxreceives an input from the core shaft and outputs drive to the fan viathe gearbox output shaft so as to drive the fan at a lower rotationalspeed than the core shaft. The fan system and the gearbox output shafttogether form an LP rotor system having a first forward whirl rotordynamic mode, 1FW. The engine has a maximum take-off speed, MTO. Aforward whirl frequency margin of:

$\frac{\begin{matrix}\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{intersection}\mspace{14mu}{of}\mspace{14mu} 1\;{FW}\mspace{14mu}{and}} \\{{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}}\end{matrix} \\{{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

is in the range from 10 to 100%.

The forward whirl frequency margin may be greater than 30%.

The forward whirl frequency margin may be greater than 20%, 30%, 40%, or50%, and/or optionally less than 100%, 90%, 80%, 70%, or 60%. Theforward whirl frequency margin may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds).

The frequency of the 1FW mode where it intersects the first engine orderline (1EO) may be calculated or read off a Campbell Diagram. Similarly,the frequency of 1EO at the MTO speed may be calculated or read off aCampbell Diagram, where 1EO intersects the MTO line.

The intersection of 1FW with the synchronous (first engine order) line,1EO, on a Campbell Diagram is commonly referred to as “synchronous 1FW”.The equation above may therefore be re-written as:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{synchronous}\mspace{14mu} 1\;{FW}\mspace{14mu}{and}} \\{{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

If the rotor first forward whirl mode (1FW) has an insufficientfrequency margin above the maximum fan speed (MTO speed), this mode canbe excited by unbalance on the rotor. The forward whirl frequency marginmay therefore be tuned appropriately, selecting a value falling within aclaimed range, to reduce or avoid excitation of this mode.

The frequency difference between synchronous 1FW and the first engineorder line at MTO speed may be in the range from 8 Hz to 45 Hz,optionally in the range from 20 Hz to 40 Hz.

The MTO speed may be in the range from 25 Hz to 45 Hz.

The MTO speed may be in the range from 25 Hz to 30 Hz and the fan mayhave a fan diameter greater than 216 cm (85 inches).

The MTO speed may be in the range from 35 Hz to 45 Hz and the fan mayhave a fan diameter less than 216 cm (85 inches).

The fan system may have a reverse travelling wave first flap mode, FanRTW, the LP rotor system may have a first reverse whirl rotor dynamicmode, Rotor RW and a backward whirl frequency margin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 15 to 50%.

The backward whirl frequency margin may be greater than 25%.

The backward whirl frequency margin may be greater than 20%, 25%, 30%,or 35%, and/or optionally less than 50%, 45%, or 40%. The backward whirlfrequency margin may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds).

The lowest frequency of either mode Fan RTW or Rotor RW at the MTO speedmay be in the range from 4 Hz to 22 Hz, optionally in the range from 5Hz to 15 Hz.

A mutual frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{matrix}( {{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}}\mspace{14mu}  \\ {{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} + {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}} )\end{matrix}}$

may be in the range from 5 to 50%.

The mutual frequency margin may be greater than 10%.

The mutual frequency margin may be greater than 10%, 15%, 20%, or 25%,and/or optionally less than 50%, 45%, 40%, or 35%. The mutual frequencymargin may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).

The frequency difference between mode Fan RTW and mode Rotor RW at theMTO speed may be in the range from 2 Hz to 15 Hz, optionally from 5 Hzto 15 Hz.

The engine may comprise a front engine structure arranged to support thefan shaft. The front engine structure may have a front engine structurenodding mode, mode FSN, which may comprise a pair of modes at similar,but not equal, natural frequencies in orthogonal directions. A frontengine structure frequency margin defined as:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}}\mspace{14mu}} \\{{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}}\mspace{14mu}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

may be in the range from 5 to 50%.

The front engine structure frequency margin may be greater than 10%.

The front engine structure frequency margin may be greater than 10%,15%, 20%, or 25%, and/or optionally less than 50%, 45%, 40%, or 35%. Thefront engine structure frequency margin may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The frequency difference between mode FSN and the highest frequency ofeither synchronous Fan RTW or synchronous Rotor RW may be in the rangefrom 2 Hz to 15 Hz, optionally from 2 Hz to 10 Hz.

The lowest natural frequency of the front structure nodding pair ofmodes may be in the range from 14 Hz to 26 Hz, optionally from 15 Hz to25 Hz.

As mentioned above with respect to 1FW, “synchronous” Fan RTW or RotorRW refers to the intersection of the respective mode (Fan RTW or RotorRW) with the first engine order line—i.e. the frequency value at whichthe lines intersect is used. Of Fan RTW and Rotor RW, whichever mode hasthe highest synchronous frequency is selected for use in the ratio shownabove.

It will be appreciated that mode FSN generally has a constant frequencyin many embodiments. Where there is any variation, the synchronousfrequency value is used (i.e. the frequency value at which the mode FSNline intersects the first engine order line).

According to a second aspect, there is provided a method of operation ofa gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan system comprising a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades and a fan shaft; and a gearbox; and a gearbox output shaftarranged to couple an output of the gearbox to the fan shaft, whereinthe gearbox receives an input from the core shaft and outputs drive tothe fan via the gearbox output shaft so as to drive the fan at a lowerrotational speed than the core shaft, and wherein the fan system and thegearbox output shaft together form an LP rotor system having a firstforward whirl rotor dynamic mode, 1FW.

The method comprises operating the engine such that a forward whirlfrequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{intersection}\mspace{14mu}{of}}\mspace{14mu}} \\{{1{FW}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{and}\mspace{14mu}{the}}\mspace{14mu}} \\{{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

is in the range from 10 to 100%.

The forward whirl frequency margin may be greater than 20%, 30%, 40%, or50%, and/or optionally less than 90%, 80%, 70%, or 60%. The forwardwhirl frequency margin may be in an inclusive range bounded by any twoof the values in the previous sentence (i.e. the values may form upperor lower bounds).

The method may comprise operating the engine at speeds up to a maximumtake-off, MTO, speed of the engine. The method may comprise operatingthe engine at the MTO speed.

The method may comprise operating the engine such that the frequencydifference between synchronous 1FW and the first engine order line atMTO speed is in the range from 8 Hz to 45 Hz, optionally from 20 Hz to40 Hz.

The fan system may have a reverse travelling wave first flap mode, FanRTW and the fan system and the gearbox output shaft together may form anLP rotor system having a first reverse whirl rotor dynamic mode, RotorRW. The method may comprise operating the engine such that a backwardwhirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}}\mspace{14mu}} \\{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{maximum}\mspace{14mu}{take}\mspace{14mu}{off}},\;{MTO},{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

is in the range from 15 to 50%.

The backward whirl frequency margin may be greater than 20%, 25%, 30%,or 35%, and/or optionally less than 50%, 45%, or 40%. The backward whirlfrequency margin may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds).

The method may comprise operating the engine such that the lowestfrequency of either mode Fan RTW or Rotor RW at the MTO speed is in therange from 4 Hz to 22 Hz, optionally from 5 Hz to 15 Hz.

The method may comprise operating the engine such that a mutualfrequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{14mu}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{matrix}( {{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{11mu}  \\ {{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} + {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}} )\end{matrix}}$

is in the range from 5 to 50%.

The method may comprise operating the engine such that the mutualfrequency margin may be greater than 10%, 15%, 20%, or 25%, and/oroptionally less than 45%, 40%, or 35%. The mutual frequency margin maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds).

The method may comprise operating the engine such that the frequencydifference between mode Fan RTW and mode RW at the MTO speed is in therange from 2 Hz to 15 Hz, optionally from 5 Hz to 15 Hz.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; and a fansystem having a reverse travelling wave first flap mode, Fan RTW. Thefan system comprises a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a fan shaft. The enginefurther comprises a gearbox, and a gearbox output shaft arranged tocouple an output of the gearbox to the fan shaft. The gearbox receivesan input from the core shaft and outputs drive to the fan via thegearbox output shaft so as to drive the fan at a lower rotational speedthan the core shaft. The fan system and the gearbox output shafttogether form a low pressure (LP) rotor system having a first reversewhirl rotor dynamic mode, Rotor RW. The engine has a maximum take-offspeed, MTO. A backward whirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}}\mspace{14mu}} \\{{RTW}\mspace{14mu}{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

is in the range from 15 to 50%.

The backward whirl frequency margin may be greater than 25%.

The backward whirl frequency margin may be greater than 20%, 25%, 30%,or 35%, and/or optionally less than 50%, 45%, or 40%. The backward whirlfrequency margin may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds).

If the rotor first reverse whirl mode (Rotor RW) and/or the reversetravelling wave first fan blade flap mode (Fan RTW) have an insufficientfrequency margin above the maximum fan shaft rotation speed (MTO,defined herein in terms of a rotation frequency of the shaft), either orboth of these modes can be excited by a forcing load that is static inthe inertial reference frame (as viewed by an outside observer viewingthe engine). Maintaining the backward whirl frequency margin within aclaimed range may therefore allow this response amplification to bereduced or avoided.

The lowest frequency of either mode Fan RTW or Rotor RW at the MTO speedmay be in the range from 4 Hz to 22 Hz, optionally in the range from 5Hz to 15 Hz.

The MTO speed may be in the range from 25 Hz to 45 Hz.

The MTO speed may be in the range from 25 Hz to 30 Hz and the fan mayhave a fan diameter greater than 216 cm (85 inches).

The MTO speed may be in the range from 35 Hz to 45 Hz and the fan mayhave a fan diameter less than 216 cm (85 inches).

The fan system may have a reverse travelling wave first flap mode, FanRTW. The LP rotor system may have a first reverse whirl rotor dynamicmode, Rotor RW.

A backward whirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}}\mspace{14mu}} \\{{RTW}\mspace{14mu}{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 15 to 50%.

A mutual frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{14mu}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{matrix}( {{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{11mu}  \\ {{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} + {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}} )\end{matrix}}$

may be in the range from 5 to 50%.

The engine may comprise a front engine structure arranged to support thefan shaft. The front engine structure may have a front engine structurenodding mode, which may comprise a pair of modes at similar, but notequal, natural frequencies in orthogonal directions. A front enginestructure frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}}\mspace{14mu}} \\{{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}}\mspace{14mu}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

may be in the range from 5 to 50%.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; and a fansystem having a reverse travelling wave first flap mode, Fan RTW. Thefan system comprises a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a fan shaft. The enginefurther comprises a gearbox, and a gearbox output shaft arranged tocouple an output of the gearbox to the fan shaft. The gearbox receivesan input from the core shaft and outputs drive to the fan via thegearbox output shaft so as to drive the fan at a lower rotational speedthan the core shaft. The fan system and the gearbox output shafttogether form an LP rotor system having a first reverse whirl rotordynamic mode, Rotor RW. The engine has a maximum take-off speed, MTO. Amutual frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{14mu}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{matrix}( {{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}}\mspace{11mu}  \\ {{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} + {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}} )\end{matrix}}$

is in the range from 5 to 50%.

The mutual frequency margin may be greater than 10%.

The mutual frequency margin may be greater than 10%, 15%, 20%, or 25%,and/or optionally less than 45%, 40%, or 35%. The mutual frequencymargin may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).

If the rotor first reverse whirl mode (Rotor RW) and reverse travellingwave first fan blade flap mode (Fan RTW) have an insufficient mutualfrequency margin (i.e. if they are too close to each other infrequency), these modes can interact such that any forcing as describedabove may excite both of these modes instead of just one.

This may again lead to deleterious increased amplitudes of vibrationalresponses. The mutual frequency margin may therefore be tunedappropriately, selecting a value falling within a claimed range, toreduce or avoid this interaction.

A backward whirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 15 to 50%.

The LP rotor system may have a first forward whirl rotor dynamic mode,1FW. A forward whirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{synchronous}\mspace{14mu} 1{FW}}\mspace{14mu}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 10 to 100%.

The engine may comprise a front engine structure arranged to support thefan shaft. The front engine structure may have a front engine structurenodding mode, which may comprise a pair of modes. The modes of the pairmay be at similar, but not equal, natural frequencies in orthogonaldirections. A front engine structure frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}}\mspace{14mu}} \\{{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}}\mspace{14mu}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

may be in the range from 5 to 50%.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; and a fansystem having a reverse travelling wave first flap mode, Fan RTW. Thefan system comprises: a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a fan shaft. The enginefurther comprises a front engine structure arranged to support the fanshaft, the front engine structure having a front engine structurenodding mode comprising a pair of modes at similar, but not equal,natural frequencies in orthogonal directions. The engine furthercomprises a gearbox, and a gearbox output shaft arranged to couple anoutput of the gearbox to the fan shaft. The gearbox receives an inputfrom the core shaft and outputs drive to the fan via the gearbox outputshaft so as to drive the fan at a lower rotational speed than the coreshaft. The fan system and the gearbox output shaft together form an LProtor system having a first reverse whirl rotor dynamic mode, Rotor RW.The engine has a maximum take-off speed, MTO. A front engine structurefrequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}}\mspace{14mu}} \\{{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}}\mspace{14mu}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

is in the range from 5 to 50%.

The front engine structure frequency margin may be greater than 10%.

The front engine structure frequency margin may be greater than 10%,15%, 20%, or 25%, and/or optionally less than 45%, 40%, or 35%. Thefront engine structure frequency margin may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

This combination of orthogonal modes may cause the front enginestructure vibration response to rotor unbalance to be elliptical inorbit. The elliptical orbit may comprise both forward and reversetravelling wave components; a mechanism is therefore presented to excitereverse whirl modes Fan RTW or Rotor RW if they are coincident ornear-coincident with the front engine structure nodding (FSN) frequency.This combined effect could rapidly increase the vibration responseamplitude to nuisance levels, or in extreme cases to potentiallydamaging/hazardous levels. The front engine structure frequency marginmay therefore be tuned appropriately, selecting a value falling within aclaimed range, to reduce or avoid this amplification mechanism.

The engine may be arranged to be mounted within a nacelle with a mass of1000 kg to 3000 kg, and optionally of 1500 kg to 2500 kg. The mass ofthe nacelle may be selected or adjusted to tune the FSN frequency.

A backward whirl frequency margin of:

$\frac{\begin{matrix}{{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}}\mspace{14mu}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 15 to 50%.

The LP rotor system may have a first forward whirl rotor dynamic mode,1FW. A forward whirl frequency margin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{synchronous}\mspace{14mu} 1{FW}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}\mspace{14mu}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 10 to 100%.

A mutual frequency margin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{pmatrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\; + \;{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\;}\end{pmatrix}}$

may be in the range from 5 to 50%.

In any of the preceding aspects, one or more of the following featuresmay apply:

-   -   A diameter of the fan may be in the range from 215 to 420 cm.    -   A diameter of the fan may be greater than or equal to 250 cm.    -   A mass of the fan may be in the range from 300 to 1000 kg.    -   A moment of inertia of the fan about the engine axis may be in        the range from 100 to 600 kg·m².    -   The tilt stiffness of the fan shaft may be in the range from        5×10⁹ to 12×10⁹N·mm/rad.    -   The radial bending stiffness of the front engine structure may        be in the range from 80 to 180 kN/mm.

The engine may comprise a front engine structure arranged to support thefan shaft. A front engine structure cantilever distance defined as thedistance between a forwardmost fan shaft bearing mounted on the frontengine structure and a radial plane at the axial position along thefront engine structure at which a front mount for the engine is locatedmay be in the range from 800 to 1700 mm.

The fan shaft may be supported by two bearings. A length of the fanshaft between the bearings may be in the range from 900 to 1800 mm.

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may further comprise a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

According to an aspect, there is provided a method of operation of a gasturbine engine for an aircraft, the engine having a maximum take-off MTOspeed, and comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan system having a reverse travelling wave first flap mode, Fan RTW,and comprising a fan located upstream of the engine core, the fancomprising a plurality of fan blades and a fan shaft; and

a gearbox and a gearbox output shaft arranged to couple an output of thegearbox to the fan shaft (36), wherein the gearbox receives an inputfrom the core shaft and outputs drive to the fan via the gearbox outputshaft so as to drive the fan at a lower rotational speed than the coreshaft; and wherein the fan system and the gearbox output shaft togetherform an LP rotor system having a first reverse whirl rotor dynamic mode,Rotor RW.

The method comprises operating the engine such that a backward whirlfrequency margin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}\mspace{14mu}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

is in the range from 15 to 50%.

The backward whirl frequency margin may be greater than 20%, 25%, 30%,or 35%, and/or optionally less than 50%, 45%, or 40%. The backward whirlfrequency margin may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds).

The method may comprise operating the engine at speeds up to a maximumtake-off speed, MTO, of the engine. The method may comprise operatingthe engine at the MTO speed.

According to an aspect, there is provided a method of operation of a gasturbine engine for an aircraft comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan system having a reverse travelling wave first flapmode, Fan RTW, and comprising a fan located upstream of the engine core,the fan comprising a plurality of fan blades and a fan shaft; a gearboxand a gearbox output shaft arranged to couple an output of the gearboxto the fan shaft, wherein the gearbox receives an input from the coreshaft and outputs drive to the fan via the gearbox output shaft so as todrive the fan at a lower rotational speed than the core shaft; whereinthe fan system and the gearbox output shaft together form an LP rotorsystem having a first reverse whirl rotor dynamic mode, Rotor RW.

The method comprises operating the engine such that a mutual frequencymargin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{pmatrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\; + \;{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\;}\end{pmatrix}}$

is in the range from 5 to 50%.

The mutual frequency margin may be greater than 10%, 15%, 20%, or 25%,and/or optionally less than 45%, 40%, or 35%. The mutual frequencymargin may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).

The method may comprise operating the engine at speeds up to a maximumtake-off, MTO, speed of the engine. The method may comprise operatingthe engine at the MTO speed.

According to a aspect, there is provided a method of operation of a gasturbine engine for an aircraft comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan system having a reverse travelling wave first flapmode, Fan RTW, and comprising a fan located upstream of the engine core,the fan comprising a plurality of fan blades and a fan shaft; a frontengine structure arranged to support the fan shaft, the front enginestructure having a front engine structure nodding mode comprising a pairof modes at similar, but not equal, natural frequencies in orthogonaldirections; and a gearbox and a gearbox output shaft arranged to couplean output of the gearbox to the fan shaft, wherein the gearbox receivesan input from the core shaft and outputs drive to the fan via thegearbox output shaft so as to drive the fan at a lower rotational speedthan the core shaft; wherein the fan system and the gearbox output shafttogether form an LP rotor system having a first reverse whirl rotordynamic mode, Rotor RW.

The method comprises operating the engine such that a front enginestructure frequency margin of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}} \\{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}} \\{{engine}\mspace{14mu}{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

is in the range from 5 to 50%.

The front engine structure frequency margin may be greater than 10%,15%, 20%, or 25%, and/or optionally less than 45%, 40%, or 35%. Thefront engine structure frequency margin may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The method may comprise operating the engine at speeds up to a maximumtake-off, MTO, speed of the engine. The method may comprise operatingthe engine at the MTO speed.

The engine of any of the preceding aspects may be used to perform themethod of any of the preceding aspects.

In the various aspects and embodiments described above, the definedfrequency margins may be arranged to remain within the defined rangesthroughout normal operation of an aircraft which the engine is arrangedto power.

The skilled person will appreciate that the most demanding conditionsfor engine vibration management may not occur around the maximum speeds(e.g. MTO speed), but rather in operating speed ranges around the speedsat which one or more of the FSN, Fan RTW and Rotor RW modes intersectthe 1EO line.

The inventors appreciated that a geared turbofan engine with a large fandiameter and a rotor system that is cantilevered forward of the frontengine mount introduces novel mass and stiffness characteristics andhence novel dynamic characteristics. The rotor stiffnesses and enginefront engine structure stiffnesses may therefore be tuned in order toreduce or avoid frequency coincidence between natural frequencies andtheir potential excitation sources.

As used herein, a “large” fan diameter may mean a fan diameter greaterthan 216 cm (85 inches), and optionally greater than 250 cm (100inches).

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. Thefan tip loading may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

As used herein, a maximum take-off (MTO) condition has the conventionalmeaning. Maximum take-off conditions may be defined as operating theengine at International Standard Atmosphere (ISA) sea level pressure andtemperature conditions +15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditionsfor the engine may therefore be defined as operating the engine at amaximum take-off thrust (for example maximum throttle) for the engine atInternational Standard Atmosphere (ISA) sea level pressure andtemperature +15° C. with a fan inlet velocity of 0.25 Mn. A maximumtake-off speed (MTO speed) is the rotational speed of the fan (and theattached fan shaft) under MTO conditions, and is measured in Hz (afrequency of rotation of the fan shaft).

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i e maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000m to15000m, for example in the range of from 10000m to 12000m, for examplein the range of from 10400m to 11600m (around 38000 ft), for example inthe range of from 10500m to 11500m, for example in the range of from10600m to 11400m, for example in the range of from 10700m (around 35000ft) to 11300m, for example in the range of from 10800m to 11200m, forexample in the range of from 10900m to 11100m, for example on the orderof 11000m. The cruise conditions may correspond to standard atmosphericconditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a sectional side view of a forward portion of a gas turbineengine;

FIG. 5 is a sectional side view of a forward portion of a gas turbineengine different from that shown in FIG. 4;

FIG. 6 is a Campbell diagram in the inertial reference frame,illustrating various vibrational modes;

FIG. 7 is the Campbell diagram of FIG. 6 with parameters A, B and Cmarked;

FIG. 8 is the Campbell diagram of FIG. 6 with parameters D, E and Fmarked;

FIG. 9 is a schematic diagram illustrating radial bending stiffness of ashaft;

FIG. 10 is a sectional side view of a forward portion of a gas turbineengine as shown in FIG. 4, illustrating how radial bending stiffness ofthe front engine structure is determined;

FIG. 11 is a schematic diagram illustrating tilt stiffness of a shaft;

FIG. 12 is a sectional side view of a forward portion of a gas turbineengine as shown in FIG. 4, illustrating how tilt stiffness of the fanshaft is determined;

FIG. 13 is a graph of displacement against load, illustrating an elasticregion within which stiffnesses of components may be determined;

FIG. 14 is a sectional side view of a gas turbine engine similar to thatshown in FIG. 1, but with a different fan shaft arrangement;

FIG. 15 illustrates various methods;

FIG. 16 schematically illustrates whirl modes of the fan and fan shaft(Rotor RW, 1FW, Fan FTW, Rotor FW);

FIG. 17 schematically illustrates the first nodding (bending) mode of afront engine structure (FSN); and

FIG. 18 schematically illustrates the reverse travelling wave (RTW)first flap mode of a fan system (Fan RTW).

DETAILED DESCRIPTION OF THE DISCLOSURE

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

The engine 10 is arranged to be mounted on a wing of an aircraft foruse, by means of one or more mounts 41. In the arrangements beingdescribed, the engine 10 is surrounded by a nacelle 21, which surroundsthe fan 23. In the example shown in FIG. 4 (in which figure the nacelle21 is not visible), the front engine mount 41 (i.e. the forwardmostmount connecting the engine 10 to the wing, however many mounts theremay be) may be described as a front core mount 41, as it connects thecore 11 directly to the wing. In the alternative example shown in FIG.5, the front mount 41 is a fan case front mount, instead of a core frontmount, as it connects the fan case 45 to the wing of the aircraft (thefan case 45 generally being positioned immediately within the nacelle21, around the axial location of the fan blade tips). The belowdescription may apply equally to engines 10 with core mounts 41 and/orfan case mounts 41; the example with a core mount shown in FIG. 4 ischosen for discussion below by way of example only; the disclosure isnot limited to such an arrangement.

The engine comprises a fan shaft 36 extending, in a geared engine 10,between a fan input position and a gearbox output position. In thearrangement shown in FIG. 14, the fan shaft 36 additionally extendsrearward of the gearbox output position, with the additional fan shaftlength providing options for fan shaft mounting rearward of the gearbox30. The fan shaft 36 transmits drive from the gearbox 30 to the fan 23.The fan shaft 36 may be defined as the torque transfer component thatcouples the output of the gearbox 30 to the fan input. For the purposesof defining the stiffness of the fan shaft 36, it is considered toextend between a fan input position (i.e. the axial position of theconnection of the fan 23 to the fan shaft 36) and a rear bearing b onthe fan shaft 36 as described below.

In various arrangements, the fan shaft 36 is supported by two bearings—afirst/forward bearing, a, located nearest the fan 23, and asecond/rearward bearing, b, located rearward of the first bearing, a.The bearings a, b, limit radial movement of the shaft 36, so enforcingnode positions for whirl modes of the fan shaft 36. In alternativearrangements, such as that shown in FIG. 14, the fan shaft 36 may besupported by more than two bearings—for example by three bearings. Thebearings a, b are both (or all) located rearward of the fan inputposition; the rotor system comprising the fan 23 and fan shaft 36 maytherefore be described as a cantilevered rotor system, as the fan 23 isonly supported by a fan shaft 36 which is supported rearward of theaxial location at which the fan 23 is connected to the fan shaft 36.

For the arrangements described in detail below, the fan shaft 36 extendsrearwardly through the gearbox 30, as shown in FIG. 14. The additionallength of the fan shaft 36 may serve to improve or facilitate axiallocation of the fan shaft 36. In the arrangement shown in FIG. 14, thegearbox 30 is a planetary gearbox, and the fan shaft 36 is thereforedriven by a gearbox output shaft 35 connected to the planet carrier 34.The fan shaft 36 is therefore driven by the rotation of the planetcarrier 34, and does not otherwise interact with the gearbox 30, despitepassing through it. In an engine 10 with a star gearbox 30, the fanshaft 36 would instead be driven by the ring gear 38.

The forward bearing, a, on the fan shaft 36 of this arrangement islocated near the fan 23, forward of the gearbox 30, and morespecifically near (and rearward of) the fan input position, i.e. theconnection between the fan shaft 36 and the fan 23. The forward bearing,a, is a roller bearing mounted to a static structure of the engine 10(and more specifically in the example shown, generally rigidly connectedto the static structure 24, including the fan outlet guide vane/enginestator). The rearward bearing, b, on the fan shaft 36 of thisarrangement is located rearward of the gearbox 30. The rearward bearing,b, is a location bearing, serving to axially locate the fan shaft 36.Bearing b is an inter-shaft bearing in the arrangement shown; axiallylocating the fan shaft 36 with respect to the core shaft 26. Anadditional bearing axially locates the core shaft 26 within the engine10.

In the arrangement shown in FIG. 14, a third bearing, c, is provided onthe fan shaft 36, between bearings a and b. This bearing c is a catcherbearing provided for safety. In alternative arrangements, this bearing cmay not be present. In various embodiments with more than two bearingson the fan shaft 36, the forwardmost bearing, nearest the fan 23, may betaken as bearing a and the rearmost bearing, furthest from the fan, asbearing b.

The engine 10 further comprises a front engine structure 42 and a powergearbox rear panel (PGB rear panel) 43.

The front engine structure 42 is substantially conical in shape in thearrangement shown in FIGS. 1 and 14, extending rearwardly and outwardlyfrom the forward bearing, a, towards the engine section stator 24. It isrigidly mounted on the engine stationary structure 24 (in thearrangement shown, the engine section stator 24 is structural and formsa part of the engine stationary structure—in other arrangements, theengine stationary structure 24 may not include the engine sectionstator), and provides a mounting for the forward bearing a, and, wherepresent, the intermediate bearing c. In the arrangement shown, the frontengine structure 42 extends from an axial position forward of thegearbox 30 to an axial position along the length of the gearbox 30. Thefront engine structure 42 therefore provides some support to the fan 23,and also provides sealing and containment for the power gear box chamber30, which generally contains an air/oil mist in operation. The forwardbearing, a, is mounted on (or an integral part of) the front enginestructure 42.

The PGB rear panel 43 may play a role in sealing and locating the gearbox 30; it may additionally provide a rotor dynamic function to theintermediate pressure compressor 14. The PGB rear panel 43 issubstantially conical in shape in the arrangement shown in FIGS. 1 and14, extending rearwardly and inwardly from a position near the enginesection stator 24 towards the rearward bearing, b. The PGB rear panel isrigidly mounted on the engine stationary structure 24 (in thearrangement shown, the engine section stator 24 is structural and formsa part of the engine stationary structure—in other arrangements, theengine stationary structure 24 may not include the engine sectionstator). In the arrangement shown, the PGB rear panel 43 extends from anaxial position along the length of the gearbox 30 to an axial positionrearward of the gearbox 30.

The PGB rear panel 43 therefore provides some support to the fan shaft36, via the core shaft 26, and also provides sealing and containment onthe rearward side of the power gear box chamber 30, which generallycontains an air/oil mist in operation.

The front engine structure 42 and the PGB rear panel 43 together form anenclosure around the gearbox chamber 30 a, shielding the rest of theengine 10 from the air/oil mist generally generated by the gearbox 30 inoperation. The front engine structure 42 and the PGB rear panel 43 arearranged not to rotate with the fan shaft 36, and may therefore bereferred to as parts of the static structure of the engine 10.

For ease of discussion herein:

-   -   a “fan system” is defined as comprising the fan 23 (fan blades        and hub) and the fan shaft 36; and    -   a “low pressure rotor system” (LP rotor system) is defined as        comprising all components 23, 36 of the fan system, and        additionally the gearbox output shaft 35 that drives the fan        shaft 36 (in the arrangement shown in FIG. 14, the gearbox        output shaft 35 is the carrier output shaft, as it is a        planetary gearbox 30).

Engine Vibrational Modes

FIGS. 4 and 5 each illustrate a forward portion of a geared turbineengine 10, with a relatively large-diameter fan 23, for example having afan diameter greater than or equal to 215 cm, and optionally greaterthan or equal to 250 cm. The fan 23 is located forward of the frontengine mount 41 in a cantilevered mounting arrangement (i.e. the fanshaft 36 is supported on only one side of the mounting position of thefan 23, namely rearward of the axial position at which the fan 23 isconnected to the fan shaft 36, such that the fan shaft 36 may be treatedas a cantilevered beam).

An engine 10 of this type may generally have three natural frequencies(modes) of interest that may be coincident or near-coincident infrequency. These modes are:

-   -   1) The first nodding (bending) mode of the front engine        structure 42 (FSN);    -   2) The reverse travelling wave (RTW) first flap mode of the fan        23 system (Fan RTW); and    -   3) The first reverse whirl (RW) rotor dynamic mode of the LP        rotor system (Rotor RW).

FIG. 6 provides a Campbell diagram in the inertial reference frame,showing various vibrational modes.

As discussed herein, rotational frequency values are notdirectional—frequencies are all given as absolute (positive) values,irrespective of rotation direction. Similarly, all frequency differencesare provided as positive values, with whichever frequency of the pair tobe compared has the lowest absolute value subtracted from whicheverfrequency has the highest absolute value. All of the vibrational modesdescribed are the lowest order vibrations of their respective type (thefundamental)—higher frequency harmonics may also be present, but invarious aircraft designs including those of the examples being describedthe fundamentals are of particular interest as several of these firstorder modes are near-coincident with each other and/or close to forcingfrequencies (unbalance or aerodynamic) likely to be present in use. Thenear-coincidence and/or forcing can amplify the vibrational responses.In addition, the skilled person would appreciate that, whilst higherorder vibrations of the same type have smaller amplitudes than the lowerorder vibrations and are therefore often less important from theperspective of their effect on the engine 10, they could present ahazard if forced and/or if near-coincident.

The first nodding (bending) mode of the front engine structure 42 may becalled the Front engine Structure Nodding mode, and referred to as FSN.The FSN line is shown as a dashed line in FIGS. 6 to 8.

The first nodding mode of the front engine structure 42 (FSN) isillustrated schematically in FIG. 17. The whole of the front enginestructure 42 bends, or “nods”, forward of the position of the rearbearing, b, and the front mount 41. It will be appreciated that FIG. 17(and correspondingly also FIGS. 16 and 18) are intended to demonstratethe mode-shape of the relevant mode, but that the displacement isexaggerated for clarity of demonstration.

The reverse travelling wave first flap mode of the fan 23 is an exampleof a Backward Whirl mode of the fan, and may be referred to as Fan RTW.The skilled person would appreciate that the fan 23 inherently has someflexibility, as required to exhibit Fan RTW vibrations, and maytherefore be referred to as a flexible fan 23. The Fan RTW line is shownas a solid dark grey line in FIGS. 6 to 8. The Fan RTW mode is mostlycomposed of movement of the fan blades, with only a small contributionfrom the fan shaft 36. FIG. 18 schematically illustrates the reversetravelling wave (RTW) first flap mode of a fan system 23, 36 (Fan RTW).As illustrated by the figure, the movement of the fan shaft 36 issmaller than that of the fan blades 23, and indeed is often negligible.

The first reverse whirl rotor dynamic mode of the fan shaft 36 isanother example of a Backward Whirl mode, and may be referred to asRotor RW. The Rotor RW line is shown as a dot-dashed black line in FIGS.6 to 8. The Rotor RW mode is mostly composed of bending of the fan shaft36, with some contribution from fan blade flex.

The two vibration modes described above, Fan RTW and Rotor RW, aretherefore both “backward whirl” (or “reverse whirl”) modes; i.e. thedirection of the whirl is opposite to the direction of rotation of therotor system 23, 36. In the example shown in FIG. 6, the lowestfrequency reverse whirl mode is the reverse travelling wave first flapmode (Fan RTW) of the flexible fan 23. The second lowest frequencyreverse whirl mode is the first reverse whirl rotor dynamic mode (RotorRW) of the fan shaft 36. However the opposite may occur in otherarrangements (i.e. Rotor RW may have a lower frequency than Fan RTW).

The Campbell Diagram (FIG. 6) also shows the synchronous line, 1EO,which may also be referred to as the first engine order line. Line 1EOrepresents the fan shaft speed operating line, and is shown as a solidblack line in FIGS. 6 to 8. FIG. 6 therefore illustrates coincidencebetween natural frequencies, ω_(n), of the modes FSN, Fan RTW and RotorRW, and the engine fan shaft speed (forcing frequency) Ω_(fan), at theintersections of the mode lines with line 1EO.

If the rotor first reverse whirl mode (Rotor RW) and/or the reversetravelling wave first fan blade flap mode (Fan RTW) have an insufficientfrequency margin above the maximum fan shaft rotation speed (i.e. if themode frequencies are too similar to the maximum fan shaft rotationfrequency/if there is not enough of a difference in frequency betweenthem), either or both of these modes can be excited by a forcing loadthat is static in the inertial reference frame (as viewed by an outsideobserver viewing the engine 10). Examples of such forcing includeaerodynamic loads on the fan blades 23, and fan blade tip rubs.

If the frequency margin were zero (i.e. if the mode frequency were equalto maximum fan shaft rotation frequency), the reverse travelling wave ofthe fan 23 and/or the rotor response would be stationary in the inertialreference frame, and hence a stationary aerodynamic load or fan bladetip rub could rapidly increase the response amplitude todamaging/hazardous levels.

A frequency margin, referred to as the backward whirl frequency margin,may therefore be tuned appropriately to avoid this responseamplification.

The maximum fan speed (i.e. MTO fan speed) is considered forestablishing this frequency margin because at lower rotor speeds thefirst reverse whirl mode (Rotor RW) and reverse travelling wave firstfan blade flap mode (Fan RTW) have higher frequencies in the inertialreference frame, while the rotor speed is lower. The maximum rotor speedcondition is therefore always the condition in which the lowest backwardwhirl frequency margin occurs in engines 10 as described.

A first parameter, A, is defined as the lowest frequency of either modeFan RTW or Rotor RW at the Maximum Take-Off (MTO) speed. In the exampleshown in FIG. 6, a line corresponding to the MTO speed (vertical dottedline) has been added to the Campbell Diagram for ease of determiningthis parameter. For the example shown, Fan RTW is lower than Rotor RW,and the value for the Fan RTW mode line where it intersects the MTO lineis therefore taken as the value for parameter A, as shown in FIG. 7.

A second parameter, B, is defined as being equal to the MTO speed. TheMTO speed is a rotational speed of the fan 23 and shaft 36, and istherefore defined in terms of a frequency—i.e. as a frequency ofrotation—for ease of comparison with the other frequencies describedherein.

The backward whirl frequency margin is expressed as A/B. The backwardwhirl frequency margin A/B may be maintained within the range from 15%to 50%, and preferably greater than 25%, in various arrangements.

If the rotor first reverse whirl mode (Rotor RW) and reverse travellingwave first fan blade flap mode (Fan RTW) have an insufficient mutualfrequency margin (i.e. if they are too close to each other infrequency), these modes can interact such that any forcing as describedabove may excite both of these modes instead of just one. This may againlead to deleterious increased amplitudes of vibrational responses.

A parameter, D, may be defined as the frequency difference between themodes Fan RTW and Rotor RW at MTO, as marked on FIG. 8. This is measuredas the difference in frequency between the intersection of the line forFan RTW and MTO, and the intersection of the line Rotor RW and MTO.

The mutual frequency margin may then be expressed as D/(A+B). Thefrequency margin D/(A+B) may be maintained within the range from 5% to50% and preferably greater than 10%, in various arrangements.

The front engine structure nodding mode (FSN) is a mode of a portion ofthe static structure, the static structure being the part of the engine10 arranged not to rotate relative to an aircraft or other structure onwhich the engine is mounted in use (i.e. not to rotate with any of theshafts 26, 36, fan 23 or turbines 19 in use).

The FSN mode can be directly excited by rotor unbalance such asunbalance of the fan 23 and/or fan shaft 36. The response to unbalancemay be amplified if the rotor unbalance at the forcing frequency (fanrotational speed, for example measured as a rotation frequency) iscoincident with the natural frequency of the FSN mode. The amplificationmay remain small provided that mode FSN does not have a frequencycoincident, or near coincident, with the frequency of Fan RTW or RotorRW. However, the vibration amplitude may be deleteriously increased ifthe FSN mode frequency is close to the frequency of Fan RTW or Rotor RW.

The frequency of the FSN mode depends on the stiffness of variousstructures 42, 24 which directly and/or indirectly support the fan shaft36, and in particular on the stiffness of the front engine structure 42.In various embodiments, the main stiffness path to the front mount plane(a) from the fan 23 may be up through the front engine structure 42,including the engine section stator 24.

In general, the stiffness of the front engine structure 42 may not beradially symmetrical—for example not being equal in orthogonaldirections due to a non-axisymmetric engine mount arrangement. As aresult the front engine structure nodding (FSN) mode is generallycomposed of a pair of modes at similar, but not equal (for example beingseparated by 0-10% only, e.g. by 2 Hz), natural frequencies inorthogonal directions in such examples. This combination of orthogonalmodes may cause the front engine structure vibration response to rotorunbalance to be elliptical in orbit, and therefore the rotor (fan 23 andfan shaft 36) housed in the front engine structure 42 may be forced byan elliptical orbit at its bearing supports a, b. The elliptical orbitmay comprise both forward and reverse travelling wave components; amechanism is therefore presented to excite reverse whirl modes Fan RTWor Rotor RW if they are coincident or near-coincident with the FSNfrequency. This combined effect could rapidly increase the vibrationresponse amplitude to nuisance levels, or in extreme cases topotentially damaging/hazardous levels. A front engine structurefrequency margin may therefore be tailored to avoid this amplificationmechanism.

A parameter, E, is defined as the frequency difference between mode FSNand the highest frequency mode of Fan RTW and Rotor RW at theirrespective synchronous natural frequencies, as shown in FIG. 8. In theexample shown in FIG. 8, Rotor RW is higher than Fan RTW, so thefrequency difference between the Rotor RW line where it crosses 1EO andthe FSN line is used. If Fan RTW were higher than Rotor RW, thefrequency difference between the Fan RTW line where it crosses 1EO andthe FSN line would be used.

A parameter, F, is defined as the lowest natural frequency of the frontengine structure nodding pair of modes (FSN), as shown in FIG. 8. On theCampbell Diagram, the FSN line shown is for the lowest natural frequencyof the front engine structure nodding pair of modes.

The front engine structure frequency margin is expressed as E/F. Thefront engine structure frequency margin E/F may be maintained within therange from 5% to 50%, and preferably greater than 10%, in variousarrangements.

In axisymmetric engine mount arrangements, the FSN mode may be composedof only a single mode, reducing or avoiding this excitation pathway;consideration of the front engine structure frequency margin may be lessimportant, or even unnecessary, in such arrangements.

The FSN mode may tend to move, and potentially bend, a nacelle 21 withinwhich the engine is mounted. A mass of the nacelle 21 may therefore beconsidered in tuning the front engine structure frequency margin, E/F.For example, the nacelle mass may be selected to be within the range of1000 kg to 3000 kg, and optionally 1500 kg to 2500 kg. In general, thefrequency of the FSN mode may reduce in proportion to the ratio of thenacelle 21 modal mass to the engine 10 modal mass, where the modal massis calculated as the mass that participates by way of kinetic energycontribution to the total energy in the FSN mode. For example, a gearedturbine engine 10 with a relatively large fan diameter and no nacellemay exhibit a FSN mode at 26 Hz. The same engine 10 mounted within anacelle with a mass of 1500 kg, may exhibit a FSN mode at 20 Hz. Thesame engine 10 mounted within a nacelle with a mass of 2500 kg, mayexhibit a FSN mode at 16 Hz. It will be appreciated that these valuesare provided by way of illustrative example only, and are not intendedto be limiting.

A geared turbine engine 10 of the type with a relatively large fandiameter and a rotor that is cantilevered forward of the front enginemount 41, as shown in FIGS. 4 and 5, may additionally have a naturalfrequency (mode) of interest at a higher frequency. This mode may beformed by a combination of two forward whirl (FW) modes:

1) The forward travelling wave first flap mode of the (flexible) fan 23system (Fan FTW); and

-   -   2) The first forward whirl rotor dynamic mode of the LP rotor        system (Rotor FW).

The two vibration modes described above, Fan FTW and Rotor FW, are both“forward whirl” modes; i.e. the direction of the whirl is the same asthe direction of rotation of the fan and LP rotor system 23, 36.

On the Campbell Diagram in the inertial reference frame (FIG. 6), aforward whirl mode is identified as 1FW (1st Forward Whirl), and markedwith a dot-dot-dashed line. 1FW may be described as a combined shapemode in that it has attributes of both the forward travelling wave firstfan flap mode (Fan FTW) shape as well as the first forward whirl rotordynamic mode shape of the fan shaft (Rotor FW).

FIG. 16 schematically illustrates the whirl modes of the fan 23 and fanshaft 36 (Rotor RW, 1FW, Fan FTW, Rotor FW). it will be appreciated thatthe mode shape is generally the same for forward and reverse whirlmodes, with the difference being the rotation direction of thewhirl—forward whirl modes rotate in the same direction as the shaft 36whereas reverse whirl modes rotate in the opposite direction to theshaft 36.

If the rotor first forward whirl mode (1FW) has insufficient frequencymargin above the maximum fan speed (MTO speed), this mode can be excitedby unbalance on the rotor 23, 36, for example by unbalance of the fan23. A high balance quality and/or control of the rotor dynamic responsemay be provided by the introduction of damping to prevent a highvibration response. The consequence of failing to prevent a highvibration response would be that vibrations of the rotor 23, 36 maycause nuisance, impose component life limitations, and/or requirefrequent fan trim balance operations. In some cases the responseamplitude could increase to damaging or hazardous levels.

A frequency margin, referred to as the forward whirl frequency margin,may therefore be tuned appropriately.

A parameter, C, is defined as the frequency difference between theintersection of 1FW with the synchronous (first engine order) line 1EO,and the intersection of MTO with 1EO, as shown on FIG. 7.

The forward whirl frequency margin is expressed as C/B, where B is themaximum take-off speed (MTO speed), which is defined in terms of thefrequency of rotation, as described above. The forward whirl frequencymargin C/B may be maintained within the range from 10% to 100%, andpreferably greater than 30%, in various arrangements.

To summarise, four frequency margins are defined herein:

TABLE 1 Frequency Margins Name Definition Range Backward whirl frequencymargin A/B 15% to 50%, optionally greater than 20%, 25%, 30%, or 35%,and optionally less than 45%, or 40% Forward whirl frequency margin C/B10% to 100%, optionally greater than 20%, 30%, 40%, or 50%, andoptionally less than 90%, 80%, 70%, or 60% Mutual frequency marginD/(A + B) 5% to 50%, optionally greater than 10%, 15%, 20%, or 25%, andoptionally less than 45%, 40%, or 35% Front engine structure frequencyE/F 5% to 50%, optionally greater than 10%, 15%, margin 20%, or 25%, andoptionally less than 45%, 40%, or 35%

In various arrangements, AB≥25%, CB≥30%, D/(A+B)≥10%, and E/F≥10%.

The following six parameters, easily obtainable from a Campbell Diagramas illustrated in FIGS. 6 to 8, are used to calculate the frequencymargins:

TABLE 2 parameters Name Definition Range A the lowest frequency ofeither mode Fan RTW or 4 Hz to 22 Hz, optionally 5 Hz to 15 Hz, andRotor RW at Maximum Take-Off Speed further optionally 6 Hz to 10 Hz BMaximum Take-Off (MTO) speed 25 Hz to 45 Hz, optionally 25 Hz to 30 Hz,e.g. for an engine with a large fan diameter (greater than 216 cm - 85inches), or optionally 35 Hz to 45 Hz, e.g. for an engine with a smallerfan diameter C the frequency difference between the intersection 8 Hz to45 Hz, optionally 20 Hz to 40 Hz, and of 1FW with 1EO and theintersection of MTO further optionally 25 Hz to 35 Hz with 1EO D thefrequency difference between mode Fan RTW 2 Hz to 15 Hz, optionally 5 Hzto 15 Hz, and and mode Rotor RW at MTO further optionally 5 Hz to 8 Hz Ethe frequency difference between mode FSN and 2 Hz to 15 Hz, optionally2 Hz to 10 Hz, and the highest frequency mode of Fan RTW and furtheroptionally 3 Hz to 5 Hz Rotor RW at their respective synchronous naturalfrequencies F the lowest natural frequency of the front engine 14 Hz to26 Hz, optionally 15 Hz to 25 Hz, and structure nodding pair of modes(FSN) further optionally 18 Hz to 22 Hz

All of these parameters have the units of frequency—Hz—and all thefrequency margins are therefore dimensionless.

In various arrangements, one, some, or all of the four frequency marginsdescribed may be maintained within the specified ranges. Various engineproperties may be controlled so as to adjust vibrational properties,including the following. The skilled person would appreciate that theengine 10 may be tuned so as to allow the frequency margin(s) to liewithin the specified ranges in a variety of different ways, as multipleparameters affect engine vibrational properties. The below examples ofengine properties are therefore provided by way of example only.

In particular, the inventors appreciated that tuning of the fan 23stiffness, fan shaft 36 stiffness, and/or the engine front enginestructure 42 stiffness may allow or facilitate the avoidance offrequency coincidence between natural frequencies and their potentialexcitation sources.

The fan diameter may be greater than or equal to 215 cm (85″) or 250 cm(100″), and optionally may be selected to be in the range from 215 cm to420 cm or from 250 cm to 370 cm (100″ to 145″). The same fan size may beused for both composite and metallic fan blades 23.

The fan mass (the mass of the fan 23, including the hub) may be in therange from 300 to 1000 kg.

The fan moment of inertia (the moment of inertia of the fan 23,including the hub) about the longitudinal engine axis may be in therange from 100 to 600 kg·m².

The fan shaft length, L, defined between the forward bearing a and therearward bearing b as shown in FIGS. 4 and 5, may be in the range from900 mm to 1800 mm. The fan shaft length, L, may be defined between theaxial centre-points of the bearings a, b. In arrangements with more thantwo bearings on the fan shaft 36, L may be defined between the fan shaftbearing closest to the fan 23 and the fan shaft bearing furthest fromthe fan 23.

The front engine structure cantilever distance, D_(c), defined as thedistance between the radial plane of the front mount 41 (the front mountplane) to the forward bearing, a, as shown in FIG. 4 may be in the rangefrom 800 mm to 1700 mm. The front engine structure cantilever distance,D_(c), may be defined between the axial centrepoint of the forwardbearing a, and the axial centrepoint of the front mount 41 (i.e. thefront mount plane is located at the axial centre point of the frontmount 41).

Radial Bending Stiffness

A radial bending stiffness is defined with reference to FIG. 9 in termsof the deformation of a cantilevered beam 900, which moves between afirst position 900 a and a second position 900 b on the application of aforce. As illustrated in FIG. 9, a force, F, applied at the free end ofthe beam 900 a in a direction perpendicular to the longitudinal axis ofthe beam causes a linear perpendicular deformation, δ, seen in thesecond position 900 b. The radial bending stiffness is the force appliedfor a given linear deformation i.e. F/δ. In the present application, theradial direction is relative to the rotational axis 9 of the engine 10,and so relates to the resistance to linear deformation in a radialdirection of the engine caused by a radial force. The beam 900, orequivalent cantilevered component, extends along the axis of rotation ofthe engine, the force, F, is applied perpendicular to the axis ofrotation of the engine 10, along any radial direction, and thedisplacement, δ, is measured perpendicular to the axis of rotation,along the line of action of the force, F. The radial bending stiffnessas defined herein has SI units of N/m, and may be scaled to alternativeunits such as kN/mm. In the present application, unless otherwisestated, the radial bending stiffness is taken to be a free-bodystiffness i.e. stiffness measured for a component in isolation in acantilever configuration, without other components present which mayaffect its stiffness.

The determination of the radial bending stiffness of the front enginestructure 42 is described with respect to FIG. 10. The front enginestructure 42 is considered in isolation (i.e. without the fan shaft 36and other components), and the deflection in response to a radial shearforce, F, applied to the front engine structure 42 at the axialcentrepoint of the forward bearing, a, is determined, with the enginestatic structures earthed (i.e. treated as rigid/not moving) at theradial plane of the front mount 41.

The deflection, δ, is measured in line with the applied force, F, at thecenterline of the forward bearing, a. Diagonal lines are used toindicate that the structure is held to be rigid in a radial planealigned with the front engine mount 41—the bending of the structureforward of this connection is measured.

In engines 10 with a non-axisymmetric engine mount 41 arrangement, theradial bending stiffness of the front engine structure 42 may not beequal in orthogonal directions. Measurements may therefore be taken at,or calculations performed for, multiple positions, e.g. two orthogonalpositions, and the lowest value may be provided for the radial bendingstiffness of the front engine structure 42. In the example beingdescribed, a mounting of the front engine structure 42 may provide anobvious asymmetry and measurements may therefore be taken in line withthe mount and perpendicular to the mount, for example. The loweststiffness may generally correspond to the lowest FSN frequency, whichmay be of interest for the minimum frequency separation to Fan RTW orRotor RW mode.

The front engine structure radial bending stiffness may be in the rangefrom 80 to 180 kN/mm.

Tilt Stiffness

A tilt stiffness is defined with reference to FIG. 11, which shows theresulting deformation of a cantilevered beam 900 from a first position900 a to a second position 900 b under a moment M applied at its freeend. The tilt stiffness is a measure of the resistance to rotation of apoint on the component at which a moment is applied. As can be seen inFIG. 11, an applied moment at the free end of the cantilevered beamcauses a constant curvature along the length of the beam between itsfree and fixed ends. The applied moment M causes a rotation θ of thepoint at which it is applied. The tilt stiffness as defined hereintherefore has SI units of Nm/rad., and may be scaled to alternativeunits such as N·mm/rad.

The determination of the tilt stiffness of the fan shaft 36 is describedwith respect to FIG. 12. Diagonal lines are used to indicate that thefan shaft 36 is held to be pinned at the bearings a and b—the bearingsa, b, are treated as rigid. The shaft 36 is treated as being pinned atthe bearings a,b, as this is representative of the boundary conditionswhen installed in the engine 10. In arrangements with more than twobearings on the fan shaft 36, the fan shaft 36 may be held to be pinnedat all such bearings.

The moment, M, is applied around a rotation axis oriented along a radiusof the engine 10 and at the axial position of the centre of gravity(CoG) of the fan assembly (i.e. the CoG of the fan 23, and not includingthe fan shaft 36). The rotation axis of the tilt moment, M, extends intothe page as drawn in FIG. 12. The fan assembly CoG axial position on thefan shaft 36 is generally at least approximately in line with, and oftenslightly forward of, the forward bearing, a, although the preciseposition may vary between different engine arrangements.

The change in angle, θ, is measured between the engine axis 9 and thetangent to the fan shaft 36 at the axial position of the CoG of the fanassembly (the point of application of the moment). The angulardeflection is measured in response to a point radial moment applied tothe fan shaft 36 in isolation (i.e. without the front engine structure42 or other components) at the fan centre of gravity, with bearingcentres pinned at “a” and “b”.

The fan shaft tilt stiffness may be in the range from 5×10⁹ to12×10⁹N·mm/rad.

FIG. 13 illustrates how the stiffnesses defined herein may be measured.FIG. 13 shows a plot of the displacement δ resulting from theapplication of a load L (e.g. a force, moment or torque) applied to acomponent for which the stiffness is being measured. At levels of loadfrom zero to L_(P) there is a non-linear region in which displacement iscaused by motion of the component (or relative motion of separate partsof the component) as it is loaded, rather than deformation of thecomponent; for example moving within clearance between parts. At levelsof load above L_(Q) the elastic limit of the component has been exceededand the applied load no longer causes elastic deformation—plasticdeformation or failure of the component may occur instead. Betweenpoints P and Q the applied load and resulting displacement have a linearrelationship. The stiffnesses defined herein may be determined bymeasuring the gradient of the linear region between points P and Q (withthe stiffness being the inverse of that gradient). The gradient may befound for as large a region of the linear region as possible to increasethe accuracy of the measurement by providing a larger displacement tomeasure. For example, the gradient may be found by applying a load equalto or just greater than L_(P) and equal to or just less than L_(Q).Values for L_(P) and L_(Q) may be estimated prior to testing based onmaterials characteristics so as to apply suitable loads. Although thedisplacement is referred to as δ in this description, the skilled personwould appreciate that equivalent principles may apply to a linear orangular displacement.

The stiffnesses defined herein, unless otherwise stated, are for thecorresponding component(s) when the engine is under cruise conditions.The stiffnesses generally do not vary significantly over the operatingrange of the engine; the stiffness at cruise conditions of the aircraftto which the engine is used (those cruise conditions being as definedelsewhere herein), or at MTO conditions, may therefore be the same asfor when the engine is not in use (i.e. off—at zero speed/on the bench).However, where the stiffness varies over the operating range of theengine, the stiffnesses defined herein are to be understood as beingvalues for when the engine is operating at cruise conditions.

FIG. 15 illustrates a method 1000 which may be performed, optionallyusing an engine 10 as described above. The method 1000 comprisesstarting up 1002 an engine 10 of an aircraft and reaching operatingconditions, and operating 1004 the aircraft. During operation 1004, theaircraft may operate at MTO speed for one or more time periods. One ormore of the following may apply:

(i) a backward whirl frequency margin (AB) of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}\mspace{14mu}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 15 to 50%;

(ii) a forward whirl frequency margin (CB) of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{synchronous}\mspace{14mu} 1{FW}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}\mspace{14mu}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$

may be in the range from 10 to 100%;

(iii) a mutual frequency margin (D/(A+B)) of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{pmatrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\; + \;{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\;}\end{pmatrix}}$

may be in the range from 5 to 50%; and/or

(iv) a front engine structure frequency margin (E/F) of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}} \\{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$

may be in the range from 5 to 50%.

The features as described above for the engine 10 may apply equivalentlyin the described methods 1000.

It will be understood that the invention is not limited to theembodiments above-described and that various modifications andimprovements can be made without departing from the concepts describedherein. Except where mutually exclusive, any of the features may beemployed separately or in combination with any other features and thedisclosure extends to and includes all combinations and sub-combinationsof one or more features described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan system comprising: a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and a fan shaft; and a gearbox and a gearbox output shaftarranged to couple an output of the gearbox to the fan shaft, whereinthe gearbox receives an input from the core shaft and outputs drive tothe fan via the gearbox output shaft so as to drive the fan at a lowerrotational speed than the core shaft; wherein the fan system and thegearbox output shaft together form an LP rotor system having a firstforward whirl rotor dynamic mode, 1FW; and wherein the engine has amaximum take-off speed, MTO, and a forward whirl frequency margin of:$\frac{\begin{pmatrix}{{t{he}}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{intersection}\mspace{14mu}{of}\mspace{14mu} 1{FW}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}} \\{\mspace{14mu}{{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\end{pmatrix}}{( {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} )}$is in the range from 10 to 100%.
 2. The gas turbine engine of claim 1,wherein the forward whirl frequency margin is greater than 30%.
 3. Thegas turbine engine of claim 1, wherein the forward whirl frequencymargin is less than 90%.
 4. The gas turbine engine of claim 1, whereinthe fan system has a reverse travelling wave first flap mode, Fan RTW,and the LP rotor system has a first reverse whirl rotor dynamic mode,Rotor RW, and a backward whirl frequency margin of:$\frac{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{or}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}\mspace{14mu}}{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}$is in the range from 15 to 50%.
 5. The gas turbine engine of claim 1,wherein the frequency difference between the intersection of 1FW and thefirst engine order line and the first engine order line at MTO speed maybe in the range from 8 Hz to 45 Hz.
 6. The gas turbine engine of claim4, wherein the lowest frequency of either mode Fan RTW or Rotor RW atthe MTO speed is in the range from 4 Hz to 22 Hz.
 7. The gas turbineengine of claim 1, wherein the MTO speed is in the range from 25 Hz to45 Hz.
 8. The gas turbine engine of claim 7, wherein the MTO speed is inthe range from 25 Hz to 30 Hz.
 9. The gas turbine engine of claim 8,wherein the fan has a fan diameter greater than 216 cm.
 10. The gasturbine engine of claim 7, wherein the MTO speed is in the range from 35Hz to 45 Hz.
 11. The gas turbine engine of claim 10, wherein the fan hasa fan diameter less than 216 cm.
 12. The gas turbine engine of claim 1,wherein a mutual frequency margin of: $\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{pmatrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\; + \;{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\;}\end{pmatrix}}$ is in the range from 5 to 50%
 13. The gas turbine engineof claim 12, wherein the frequency difference between mode Fan RTW andmode Rotor RW at the MTO speed is in the range from 2 Hz to 15 Hz. 14.The gas turbine engine of claim 1, wherein the engine comprises a frontengine structure arranged to support the fan shaft; and the front enginestructure has a front engine structure nodding mode comprising a pair ofmodes at similar, but not equal, natural frequencies in orthogonaldirections, and a front engine structure frequency margin of:$\frac{\begin{matrix}{a\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{FSN}\mspace{14mu}{and}} \\{{the}\mspace{14mu}{highest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{synchronous}} \\{{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}\mspace{14mu}{synchronous}\mspace{14mu}{Rotor}\mspace{14mu}{RW}}\end{matrix}}{\begin{matrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{natural}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{front}} \\{{structure}\mspace{14mu}{nodding}\mspace{14mu}{pair}\mspace{14mu}{of}\mspace{14mu}{modes}}\end{matrix}}$ is in the range from 5 to 50%.
 15. The gas turbine engineof claim 14, wherein the frequency difference between mode FSN and thehighest frequency of either synchronous Fan RTW or synchronous Rotor RWis in the range from 2 Hz to 15 Hz.
 16. The gas turbine engine of claim14, wherein the lowest natural frequency of the front structure noddingpair of modes is in the range from 14 Hz to 26 Hz.
 17. A method ofoperation of a gas turbine engine for an aircraft, the engine having amaximum take-off (MTO) speed, and comprising: an engine core comprisinga turbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan system and comprising a fan located upstream of theengine core, the fan comprising a plurality of fan blades and a fanshaft; and a gearbox and a gearbox output shaft arranged to couple anoutput of the gearbox to the fan shaft, wherein the gearbox receives aninput from the core shaft and outputs drive to the fan via the gearboxoutput shaft so as to drive the fan at a lower rotational speed than thecore shaft; wherein the fan system and the gearbox output shaft togetherform an LP rotor system having a first forward whirl rotor dynamic mode,1FW, the method comprising: operating the engine such that a forwardwhirl frequency margin of: $\frac{\begin{pmatrix}{{t{he}}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{intersection}\mspace{14mu}{of}\mspace{14mu} 1{FW}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}} \\{\mspace{14mu}{{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{engine}\mspace{14mu}{order}\mspace{14mu}{line}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\end{pmatrix}}{( {{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}} )}$is in the range from 10 to 100%.
 18. The method of claim 17, comprisingoperating the engine such that the frequency difference between theintersection of 1FW and the first engine order line and the first engineorder line at MTO speed is in the range from 8 Hz to 45 Hz.
 19. Themethod of claim 17, wherein the fan system has a reverse travelling wavefirst flap mode, Fan RTW, and the fan system and the LP rotor system hasa first reverse whirl rotor dynamic mode, Rotor RW; the methodcomprising: operating the engine such that a mutual frequency margin of:$\frac{\begin{matrix}{{the}\mspace{14mu}{frequency}\mspace{14mu}{difference}\mspace{14mu}{between}\mspace{11mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}} \\{{and}\mspace{14mu}{mode}\mspace{14mu}{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\end{matrix}}{\begin{pmatrix}{{the}\mspace{14mu}{lowest}\mspace{14mu}{frequency}\mspace{14mu}{of}\mspace{14mu}{either}\mspace{14mu}{mode}\mspace{14mu}{Fan}\mspace{14mu}{RTW}\mspace{14mu}{or}} \\{{{{Rotor}\mspace{14mu}{RW}\mspace{14mu}{at}\mspace{14mu}{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}\; + \;{{the}\mspace{14mu}{MTO}\mspace{14mu}{speed}}}\;}\end{pmatrix}}$ is in the range from 5 to 50%.
 20. The method of claim19, comprising operating the engine such that the frequency differencebetween mode Fan RTW and mode RW at the MTO speed is in the range from 2Hz to 15 Hz.